A gas turbine engine may be used to power various types of vehicles and systems, such as aircraft engines and auxiliary power units in aircraft. In a typical configuration, the turbines of such engines include rows of airfoils, such as stator vanes and rotor blades disposed in an alternating sequence along the axial length of a generally annular hot gas flow path. The rotor blades are mounted at the periphery of one or more rotor disks mounted on a main engine shaft. Hot combustion gases are delivered from an engine combustor to the annular hot gas flow path, thus resulting in rotary driving of the rotor disks and the main engine shaft to provide an engine output.
In most gas turbine engine applications, it is desirable to regulate the operating temperature of the engine components in order to prevent overheating and potential mechanical failures attributable thereto. For example, the airfoils of the turbines require such cooling, particularly at the trailing edge that is significantly hotter than the remainder of the airfoil. Elevated temperatures at the trailing edge may result in high thermal stress and oxidation, and thus limit the life of the airfoil. Accordingly, airfoils may be cooled using a source of relatively cool air, such as compressor discharge air, that flows through an internal cooling circuit within the airfoil and that exits through slots at the trailing edge. While these configurations may be effective, there remains a need for enhanced cooling of the trailing edges of turbine airfoils.
Accordingly, it is desirable to provide a gas turbine engine with turbine airfoils having improved cooling arrangements for the trailing edges. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description of the invention and the appended claims, taken in conjunction with the accompanying drawings and this background of the invention.